Gas turbine engine with improved vigv shielding

ABSTRACT

A gas turbine engine includes: a fan rotating about an engine main axis; a core duct; an engine core; an Engine Section Stator (ESS) including a plurality of ESS vanes and arranged in the core duct downstream of the fan; and a plurality of variable inlet guide vanes (VIGV) adapted to rotate about a pivot axis and arranged in the core duct downstream of the ESS. The VIGV vanes are arranged angularly rotated with respect to the ESS vanes such that the VIGVs are shielded by the ESS, thereby protecting the VIGVs from icing and from ice shedding from the ESS vanes.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. application Ser. No.17/514,977 filed Oct. 29, 2021, which claims the benefit of priorityfrom UK Patent Application Number 2018264.8 filed on Nov. 20, 2020, theentire contents of which are incorporated herein by reference.

BACKGROUND 1. Field of the Disclosure

The present disclosure relates to a gas turbine engine and moreparticularly to a gas turbine engine including components, such asvanes, which may be subject to icing.

2. Description of the Related Art

Gas turbine engines are typically employed to power aircraft. A gasturbine engine generally comprises, in axial flow series from front toaft, a fan, a core and an exhaust nozzle. The core comprises one or morecompressors, a combustor, and one or more turbines. An air intake isprovided for the engine. Air entering the air intake is accelerated bythe fan to produce two air flows: a first air flow (core engine flow)through a core duct into the compressor and a second air flow (bypassflow) to provide propulsive thrust. Air entering the compressor iscompressed, mixed with fuel and then fed into the combustor, wherecombustion of the air/fuel mixture occurs. The high temperature and highenergy exhaust fluids are then fed to the turbine, where the energy ofthe fluids is converted to mechanical energy to drive the compressor inrotation by suitable interconnecting shaft.

An Engine Section Stator (ESS) with a plurality of ESS vanes isgenerally provided at the inlet of the core duct. The ESS vanes guidethe air flow entering the core duct. The ESS vanes may be structural,i.e. the ESS may be provided to support load between an inner and anouter casing member; or non-structural. Furthermore, Inlet Guide Vanes(IGVs) may be provided in the core duct downstream of the ESS vanes andupstream of the compressor to further guide the air flow entering thecompressor inlet. IGVs may be variable (VIGVs), i.e. rotatable about aradial pivot, to adjust the airflow depending on different engineoperation conditions.

In appropriate atmospheric conditions, icing of components may occur atany time when the engine is running, that is to say in use. Thisincludes ground running, at idle or at higher engine speeds, as well asoperation in flight. In such circumstances ice may build up on ESS vanesand IGVs. In particular, there is a risk of icing when the design of theengine is such that the fan imparts only a low temperature rise to theair, and to a lesser extent the risk is exacerbated by lower fan bladerotational speed, such as in large and medium geared gas turbineengines, and a reduced number of fan blades.

Ice attached to the surface of a vane may effectively change thegeometry of the vane, such that oncoming flow is presented with asurface that is not to design specification, and add unnecessary weight.This may ultimately result in lower engine efficiency and/orperformance. Furthermore, ice formed on the vanes can break awaytherefrom in a process known as “shedding”, which can cause ice impactsthat may damage downstream components of the engine, in particularrotating components of the engine. For example ice shed from the ESSvanes and/or (V)IGVs may strike and damage rotating downstreamcompressor blade rows.

Many systems have been developed in attempts to reduce the formation ofice on the vanes or to minimise the damage caused by shedding from thevanes.

Conventional anti-icing and de-icing systems use hot air, bled from thecompressor and ducted to the areas of the engine requiring de-icing, orelectrical heating of the parts concerned; sometimes a combination ofthe two is used. Other known systems have used ducted hot oil,microwaves or chemical de-icing means.

A disadvantage of known anti-icing and de-icing systems is that theyrequire additional hardware, in the form of bleeds and ducting for hotair, or heating elements and their associated control systems, which addweight and complexity to the engine. In addition, the need for warmedand pressurised air, or for electrical power, is detrimental to theoverall efficiency and performance of the engine.

The present disclosure seeks to provide a gas turbine engine withimproved VIGV anti-icing capability that addresses the above issues andovercomes the disadvantages of the known techniques.

SUMMARY

Accordingly there is provided a gas turbine engine comprising a fanrotating about an engine main axis and generating a core airflow and abypass airflow; a core duct, across which the core airflow flows; and anengine core. The engine core comprises a compressor for compressing thecore airflow and comprising a plurality of stages, each stage comprisinga row of rotor blades and a row of stator vanes, a first stage of saidplurality of stages being arranged at an inlet of the compressor;combustion equipment; and a turbine connected to the compressor througha shaft. The gas turbine engine further comprises an Engine SectionStator (ESS) comprising a plurality of ESS vanes and arranged in thecore duct downstream of the fan, each ESS vane comprising an ESS leadingedge and an ESS trailing edge; and a plurality of variable inlet guidevanes (VIGV) adapted to rotate about a pivot axis and arranged in thecore duct downstream of the ESS and upstream of the compressor, eachvariable inlet guide vanes (VIGV) comprising a VIGV leading edge and aVIGV trailing edge. A mid-span ESS leading edge point is arranged at afirst radius from the engine main axis and a mid-span VIGV leading edgepoint is arranged at a second radius from the engine main axis, themid-span ESS leading edge point being at an axial distance L from themid-span VIGV leading edge point, wherein the ratio ΔR/L of a differenceΔR between the first radius and the second radius to the axial distanceL is comprised between 0.23 and 0.70, preferably between 0.40 and 0.70,and wherein the VIGV vanes are arranged angularly rotated with respectto the ESS vanes such that first longitudinal planes passing throughrespective 70% span ESS leading edge points are rotated with respect tocorresponding second longitudinal planes passing through respective 70%span VIGV pivot axis points by a rotation angle α comprised between 0.1°and 6°.

In embodiments, the rotation angle α may be comprised between 0.1° and5°, or between 0.1° and 4°, or between 0.1° and 3°, or between 0.2° and6°, or between 0.3° and 6°, or between 0.4° and 6°, or between 0.5° and6°, or between 0.5° and 5°.

During flight through supercooled cloud, or ground operation in freezingfog, droplets of supercooled liquid water are ingested into the enginewhere they subsequently collide with rotating or stationary surfaces andcan freeze into ice. For the engine core duct the primary accretionsites are the ESS vane leading edge and pressure surface, the VIGVs, thecore splitter, and the low pressure compressor outer annulus linebetween the core splitter and the VIGVs.

The higher inertia of the droplets prevents them from turning as quicklyas the aft around blades or vanes leading to regions immediately behindthese blades or vanes where there is very low concentration ofsupercooled liquid water droplets.

The present inventors have discovered that VIGVs may be protected fromicing by positioning the VIGVs with respect to the ESS vanes such thatthe VIGVs are located within these low droplet concentration regions(shielded by the ESS), thereby protecting the VIGVs from ice buildup.

Because the water droplets (that stick and form ice on the ESS or VIGV)are pushed radially outwards by the centrifugal force imparted by thefan, there is a concentrating effect towards the outer annulus line ofthe ESS and core inlet and therefore ice tends to form on outer span ofESS vanes and VIGVs, namely between 40% and 100% span. It is thereforeimportant to arrange the radially outer part of the VIGVs such as to beshielded by the ESS.

In the present disclosure, upstream and downstream are with respect tothe air flow through the compressor; and front and rear is with respectto the gas turbine engine, i.e. the fan being in the front and theturbine being in the rear of the engine.

As used herein, a longitudinal plane is a plane containing the enginemain axis, or, in other words, passing through the engine main axis, andextending radially therefrom. In this light, the first and secondlongitudinal planes pass through the engine main axis.

The ESS may be arranged at an inlet of the core duct. The compressor maybe arranged immediately downstream of the plurality of the VIGVs.

The axial distance L may be comprised between 80 mm and 650 mm, forexample between 80 mm and 500 mm, or between 80 mm and 400 mm, orbetween 80 mm and 300 mm, or between 80 mm and 250 mm, or between 100 mmand 400 mm, or between 100 mm and 300 mm, or between 120 mm and 400 mm,or between 120 mm and 350 mm, or between 120 mm and 300 mm, or between120 mm and 250 mm, or between 150 mm and 550 mm.

The difference ΔR may be comprised between 60 mm and 280 mm, for examplebetween 60 mm and 260 mm, or between 60 mm and 240 mm, or between 60 mmand 230 mm, or between 70 mm and 280 mm, or between 70 mm and 260 mm, orbetween 70 mm and 240 mm, or between 70 mm and 230 mm, or between 75 mmand 280 mm, or between 75 mm and 260 mm, or between 75 mm and 240 mm, orbetween 75 mm and 230 mm.

The ESS may comprise 40 to 80 ESS vanes, preferably 40 to 60.

The VIGVs are in number of between 40 and 80, preferably between 40 and60.

The gas turbine engine may comprise an equal number of ESS vanes andVIGVs.

The ESS leading edge may not extend linearly along a radial direction.

The pivot axis of the VIGVs may be substantially radial.

As the pivot axis of the VIGVs may extend along a radial direction,whereas the ESS leading edge may extend along a curved line, the anglebetween corresponding longitudinal planes passing through ESS leadingedge points and VIGV pivot axis points at different span heights mayvary along the span.

For example, at 90% span the rotation angle α1, i.e. the angle formedbetween longitudinal planes passing through respective 90% span ESSleading edge points and longitudinal planes passing through respective90% span VIGV pivot axis points, may be less or greater than therotation angle α at 70% span. For example the rotation angle α1 at 90%span may be less than the rotation angle α at 70% span and may be in therange 0.05° to 5°, for example in the range 0.2° to 5°, or in the range0.2° to 4°. In alternative embodiments, the rotation angle α1 at 90%span may be greater than the rotation angle α at 70% span and may be inthe range 0.2° to 6°, for example in the range 0.2° to 5°, or in therange 0.4° to 5°.

For example at 50% span the rotation angle α2, i.e. the angle formedbetween longitudinal planes passing through respective 50% span ESSleading edge points and longitudinal planes passing through respective50% span VIGV pivot axis points, may be less or greater than therotation angle α at 70% span. For example the rotation angle α2 at 50%span may be less than the rotation angle α at 70% span and may be in therange 0.05° to 5°. In alternative embodiments, the rotation angle α2 at50% span may be greater than the rotation angle α at 70% span and may bein the range 0.2° to 6°

The gas turbine engine may further comprise a strut arranged in the coreduct between the ESS and the VIGVs. For example, the gas turbine enginemay comprise one or more struts. The one or more struts may be arrangedin the core duct downstream of the ESS. The one or more struts may bearranged in the core duct upstream of the compressor, in particularupstream of the VIGVs. In embodiments comprising the one or more struts,the axial distance L may be comprised between 300 mm and 650 mm, forexample between 350 mm and 650 mm, or between 400 mm and 650 mm, orbetween 450 mm and 650 mm, or between 300 mm and 600 mm, or between 350mm and 600 mm, or between 400 mm and 600 mm, or between 450 mm and 600mm, or between 300 mm and 550 mm, or between 350 mm and 550 mm, orbetween 400 mm and 550 mm. Moreover, the difference ΔR may be comprisedbetween 100 mm and 280 mm, for example between 120 mm and 280 mm, orbetween 140 mm and 280 mm, or between 160 mm and 280 mm, or between 100mm and 260 mm, or between 120 mm and 260 mm, or between 140 mm and 260mm, or between 160 mm and 260 mm, or between 100 mm and 240 mm, orbetween 120 mm and 240 mm, or between 140 mm and 240 mm, or between 160mm and 240 mm.

The compressor may be a first compressor, the turbine may be a firstturbine, and the shaft may be a first shaft. The engine core may furthercomprise a second compressor downstream of the first compressor, asecond turbine upstream of the first turbine, and a second shaftconnecting the second turbine with the second compressor. Where morethan one compressor are present, the first compressor may be the mostupstream compressor in the core duct.

The disclosure may be particularly advantageous in large and mediumgeared gas turbine engines, wherein icing and ice shedding are ofparticular concern.

Accordingly, the gas turbine engine may further comprise a reductiongearbox that receives an input from the shaft and outputs drive to thefan so as to drive the fan at a lower rotational speed than the shaft.

The gearbox may have a reduction ratio comprised between 3.1 and 3.8,preferably between 3.1 and 3.7, more preferably between 3.2 and 3.6.

The gas turbine engine may feature a fan with a diameter comprisedbetween 240 cm and 400 cm, preferably between 240 cm and 380 cm,preferably between 300 cm and 390 cm, more preferably between 330 cm and380 cm, even more preferably between 335 cm and 360 cm.

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan isa lower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any desired reduction ratio (defined as the rotationalspeed of the input shaft divided by the rotational speed of the outputshaft), for example greater than 2.5, for example in the range of from 3to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1,3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratiomay be, for example, between any two of the values in the previoussentence. Purely by way of example, the gearbox may be a “star” gearboxhaving a ratio in the range of from 3.1 or 3.2 to 3.8. In somearrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38, 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.32. These ratios maycommonly be referred to as the hub-to-tip ratio. The radius at the huband the radius at the tip may both be measured at the leading edge (oraxially forwardmost) part of the blade. The hub-to-tip ratio refers, ofcourse, to the gas-washed portion of the fan blade, i.e. the portionradially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches),260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm(around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160inches) or 420 cm (around 165 inches). The fan diameter may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds), for examplein the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cmto 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for examplein the range of from 1800 rpm to 2300 rpm, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 330 cm to 380 cm maybe in the range of from 1200 rpm to 2000 rpm, for example in the rangeof from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpmto 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (allvalues being dimensionless). The fan tip loading may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds), for example in the range offrom 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratiomay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypassduct may be substantially annular. The bypass duct may be radiallyoutside the core engine. The radially outer surface of the bypass ductmay be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹ s, 105 Nkg⁻¹ s, 100 Nkg⁻¹ s, 95 Nkg⁻¹ s, 90 Nkg⁻¹ s, 85 Nkg⁻¹ s or80 Nkg⁻¹ s. The specific thrust may be in an inclusive range bounded byany two of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 80 Nkg⁻¹ s to100 Nkg⁻¹ s, or 85 Nkg⁻¹ s to 95 Nkg⁻¹ s. Such engines may beparticularly efficient in comparison with conventional gas turbineengines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). Purely by way of example, agas turbine as described and/or claimed herein may be capable ofproducing a maximum thrust in the range of from 330 kN to 420 kN, forexample 350 kN to 400 kN. The thrust referred to above may be themaximum net thrust at standard atmospheric conditions at sea level plus15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),with the engine static.

In use, the temperature of the flow at he entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane, At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e, thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e, the values may form upper or lowerbounds), for example in the range of from 1800K to 1950K. The maximumTET may occur, for example, at a high thrust condition, for example at amaximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of, the gas turbine engine that providesa thrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 is a longitudinal section of a core duct;

FIG. 5 is a schematic front view of the core duct taken along arrows 6a-6 a and 6 b-6 b of FIG. 4 to show ESS vanes and VIGVs with respectivefirst and second longitudinal planes and corresponding rotation angle;and

FIG. 6 is a schematic isometric view of the core duct with an ESS vaneand a corresponding VIGV, showing a first longitudinal plane passingthrough a 70% span VIGV pivot axis point and a corresponding secondlongitudinal plane passing through a respective 70% span ESS leadingedge point.

DETAILED DESCRIPTION OF THE DISCLOSURE

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis, or engine main axis 9. The engine 10 comprises an air intake 12and a propulsive fan 23 that generates two airflows: a core airflow Aand a bypass airflow B. The gas turbine engine 10 comprises a core 11that receives the core airflow A. The engine core 11 comprises, in axialflow series, a low pressure compressor 14, a high-pressure compressor15, combustion equipment 16, a high-pressure turbine 17, a low pressureturbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gasturbine engine 10 and defines a bypass duct 22 and a bypass exhaustnozzle 18. The bypass airflow B flows through the bypass duct 22. Thefan 23 is attached to and driven by the low pressure turbine 19 via ashaft 26 and an epicyclic gearbox 30. In some arrangements, the gasturbine engine 10 may not comprise a gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2 . The low pressure turbine 19 (see FIG. 1 ) drives the shaft26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclicgear arrangement 30. Radially outwardly of the sun gear 28 andintermeshing therewith is a plurality of planet gears 32 that arecoupled together by a planet carrier 34. The planet carrier 34constrains the planet gears 32 to precess around the sun gear 28 insynchronicity whilst enabling each planet gear 32 to rotate about itsown axis. The planet carrier 34 is coupled via linkages 36 to the fan 23in order to drive its rotation about the engine axis 9. Radiallyoutwardly of the planet gears 32 and intermeshing therewith is anannulus or ring gear 38 that is coupled, via linkages 40, to astationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3 . Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3 . There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2 . For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2 .

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1 ), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

FIG. 4 illustrates an inlet duct 50 through which the core flow A entersthe core 11 of the gas turbine engine 10. The inlet duct 50 is definedbetween an inner wall 64 and an outer wall 66 provided radiallyoutwardly of the inner wall 64. An annular spacing between a radiallyouter surface 65 of the inner wall 64 and a radially inner surface 67 ofthe outer wall 66 defines the core duct which contains the core airflowA.

Arranged in the inlet duct 50, there are in flow series an EngineSection Stator (ESS) 52 with a plurality of vanes 54, a row of VariableInlet Guide Vanes (VIGVs) 58, and a first rotor 60 of the low pressurecompressor 14 with a plurality of first blades 62. Optionally, a strut56 may be arranged in the inlet duct 50 between the ESS and the firstrotor 60 of the low pressure compressor 14, for example between the ESSand the VIGVs 58. The strut 56 may be omitted in case the ESS isstructural, i.e. if the ESS 52 is provided to support load between theinner wall 64 and the outer wall 66.

In general, the inner wall 64 and the outer wall 66 feature a curvedprofile and extend inwardly from the ESS 52 to the VIGVs 58 towards theengine main axis 9.

The ESS vanes 54 and the VIGVs direct the air entering the inlet duct 50appropriately towards the compressor inlet, for example to improve theengine performance and avoid flow separation at the first blades 62.

The ESS vanes 54 are uniformly spaced circumferentially around theengine main axis 9 and have an aerofoil profile with an ESS leading edge72 and an ESS trailing edge 74. The aerofoil profile extends in achordwise direction from the leading edge to the trailing edge and in aspanwise direction from a first, radially inward end to a second,radially outward end. In this light, the ESS leading edge 72 extendsfrom a first, radially inward, point 71 corresponding to 0% span to asecond, radially outward, point 73 corresponding to 100% span. Point 71may be referred to as 0% span ESS leading edge point 71 and point 73 as100% span ESS leading edge point 73.

A mid-span ESS leading edge point PLE1 is defined on the ESS leadingedge 72 at 50% span, midway between the inner wall 64 and the outer wall66, or in other words midway between the 0% span ESS leading edge point71 and the 100% span ESS leading edge point 73. The mid-span ESS leadingedge point PLE1 is arranged at a first radius R1 from the engine mainaxis 9.

The VIGVs 58 are uniformly arranged circumferentially about the enginemain axis 9 immediately upstream of the first rotor 60 of the lowpressure compressor 14. The VIGVs 58 are rotatable about respectiveradial, or nearly radial, pivot axes 80, for example by means of arotation mechanism per se known and therefore not illustrated in detail.In a typical arrangement each individual vane in a VIGV row is typicallysupported in two journal bearings at the radially inner and outer endsof the vane aerofoil section. The journal bearings permit the vaneaerofoil to rotate or pivot about its spanwise axis. This axis istypically radial, or nearly radial, relative to the compressor or enginemain axis 9.

The VIGVs 58 have an VIGV leading edge 76 and a VIGV trailing edge 78and extend in a chordwise direction from the VIGV leading edge 76 to theVIGV trailing edge 78 and in a spanwise direction from a first, radiallyinward, end to a second, radially outward, end. In this light, the VIGVleading edge 76 extends from a first, radially inward point 75corresponding to 0% span to a second, radially outward point 77corresponding to 100% span. Point 75 may be referred to as 0% span VIGVleading edge point 75 and point 77 as 100% span VIGV leading edge point77.

A mid-span VIGV leading edge point PLE2 is defined on the VIGV leadingedge 76 at 50% span, midway between the inner wall 64 and the outer wall66, or in other words midway between the 0% span VIGV leading edge point75 and 100% span VIGV leading edge point 77. The mid-span VIGV leadingedge point PLE2 is arranged at a second radius R2 from the engine mainaxis 9. The first radius R1 is generally greater than the second radiusR2 and the difference ΔR between R1 and R2 is comprised between 60 mmand 280 mm, preferably between 150 mm and 260 mm. In an embodiment, thedifference ΔR between R1 and R2 is for example 200 mm.

The mid-span ESS leading edge point PLE1 is axially distanced from themid-span VIGV leading edge point PLE2 by an axial distance L. Forexample, the axial distance L is comprised between 300 mm and 650 mm,preferably between 450 mm and 650 mm. In an embodiment, the axialdistance L is for example 500 mm. The ratio ΔR/L of the difference ΔRbetween the first radius R1 and the second radius R2 to the axialdistance L is comprised between 0.23 and 0.70, preferably between 0.40and 0.70. In an embodiment, the ratio ΔR/L is for example equal to 0.45.

The VIGVs 58 are arranged angularly rotated about the engine main axes 9by a rotation angle α with respect to the ESS vanes 54, so that theVIGVs 58 are positioned in a shielded position with respect to the ESSvanes 54 to reduce water droplets contacting the VIGVs and therefore icebuildup. In particular, the VIGVs 58 are positioned such that a radiallyouter part of the VIGVs 58 is shielded. The radially outer part may bedefined as the part of the VIGVs 58 between 40 and 100% span height.

The mutual arrangement of the ESS vanes 54 and VIGVs 58 will be furtherdescribed with reference to FIGS. 5 and 6 .

FIG. 5 is a schematic, simplified front view of the inlet duct 50 takenalong arrows 6 a-6 a and 6 b-6 b of FIG. 4 to show ESS vanes 54 andcorresponding VIGVs 58. In substance, FIG. 5 shows a first cross-section6 a-6 a taken at a first transversal plane passing through points 90defined at 70% span on the ESS leading edge 72 and a secondcross-section 6 b-6 b taken at a second transversal plane passingthrough points 92 defined at 70% span on the pivot axes 80. Points 90may be referred to as 70% span ESS leading edge points 90 and points 92as 70% span VIGV pivot axis points 92.

For sake of simplicity, not all of the ESS vanes 54 and VIGVs 58 of thecore duct 50 have been illustrated.

FIG. 5 further illustrates first longitudinal planes LP1 passing throughrespective 70% span ESS leading edge points 90 and second longitudinalplanes LP2 passing through respective 70% span VIGV pivot axis points92.

The first longitudinal plane LP1 and second longitudinal plane LP2 passthrough the engine main axis 9.

Each first longitudinal plane LP1 is angularly rotated with respect to acorresponding second longitudinal plane LP2 by the rotation angle α.

For each first longitudinal plane LP1 the corresponding secondlongitudinal plane LP2 is defined as the adjacent second longitudinalplane LP2 in a clockwise direction as seen from the front of the engine,namely as shown in FIG. 5 . Similarly, for each ESS vane 54 acorresponding VIGV 58 is defined as the adjacent VIGV 58 in a clockwisedirection as seen from the front of the engine.

The ESS vanes 54 and VIGVs 58 are in a same number comprised between 40and 80, for example 48, or 54, or 60, and, as both the ESS vanes 54 andVIGVs 58 are uniformly angularly arranged about the engine main axis 9,for each pair of ESS vane 58 and corresponding VIGV 58 the rotationangle α is the same.

It has to be noted that the VIGVs 58 are rotatable about the respectivepivot axes 80 and therefore the VIGV leading edge 76 moves with respectto the ESS leading edge 72, whereas the pivot axis 80 does not.Consequently, the mutual position of the ESS leading edge 72 and thepivot axis 80, for any span height, does not change, even if the VIGVrotates. In other words, the rotation angle α between the firstlongitudinal plane LP1 and the second longitudinal plane LP2 does notdepend on the VIGV rotation and does not vary with the VIGV rotationabout the pivot axis 80.

Furthermore, it has to be noted that the rotation angle betweencorresponding longitudinal planes passing through points at differentspan heights (i.e. at span heights different from 70%) on the ESSleading edge 72 and on the VIGV pivot axis 80 may vary with span heightbecause of the generally 3D shape of the ESS vanes, and/or thenon-perfectly radially orientation of the pivot axis 80, but not withVIGV rotation.

FIG. 6 shows an isometric view of a detail of an embodiment of the coreduct 50 with an ESS vane 54 and a corresponding VIGV 58, arrangedbetween the inner wall 64 and the outer wall 66. The embodiment of FIG.6 does not comprise the strut 56.

FIG. 6 further shows the first longitudinal plane LP1 passing throughthe 70% span ESS leading edge point 90 of the ESS vane 54 and the secondlongitudinal plane LP2 passing through the 70% span VIGV pivot axispoint 92 of the VIGV 58. The second longitudinal plane LP2 is rotated bythe rotation angle α from the first longitudinal plane LP1 about theengine main axis 9 in a clockwise direction when seen from the front ofthe engine. In embodiments the rotation angle α is comprised between0.1° and 6°, for example 2°, or 2.5°, or 3°.

The first longitudinal plane LP1 intersects the inner wall 64 along afirst line 94 and the outer wall 66 along a second line 96. Both firstand second lines 94, 96 are shown as dotted lines in FIG. 6 .

The second longitudinal plane LP2 intersects the inner wall 64 along athird line 98 and the outer wall 66 along a fourth line 100. Both thethird and fourth lines 98, 100 are shown as dotted lines in FIG. 6 .

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the scope of the invention asdescribed in the appended claims. Except where mutually exclusive, anyof the features may be employed separately or in combination with anyother features and the disclosure extends to and includes allcombinations and sub-combinations of one or more features describedherein.

We claim:
 1. A gas turbine engine comprising: a fan rotating about anengine main axis and generating a core airflow and a bypass airflow; acore duct, across which the core airflow flows; an engine corecomprising: a compressor for compressing the core airflow and comprisinga plurality of stages, each stage comprising a row of rotor blades and arow of stator vanes, a first stage of said plurality of stages beingarranged at an inlet of the compressor; combustion equipment; and aturbine connected to the compressor through a shaft; an Engine SectionStator (ESS) comprising a plurality of ESS vanes and arranged in thecore duct downstream of the fan, each ESS vane comprising an ESS leadingedge and an ESS trailing edge; and a plurality of variable inlet guidevanes (VIGVs) adapted to rotate about a pivot axis and arranged in thecore duct downstream of the ESS and upstream of the compressor, eachvariable inlet guide vanes (VIGVs) comprising a VIGV leading edge and aVIGV trailing edge, and wherein the plurality of VIGVs are arrangedangularly rotated with respect to the plurality of ESS vanes such thatfirst longitudinal planes passing through respective 70% span ESSleading edge points are angularly rotated with respect to correspondingsecond longitudinal planes passing through respective 70% span VIGVpivot axis points by a rotation angle α comprised between 0.1° and 6°.2. The gas turbine engine of claim 1, further comprising a reductiongearbox that receives an input from the shaft and outputs drive to thefan so as to drive the fan at a lower rotational speed than the shaft.3. The gas turbine engine of claim 2, wherein the gearbox has areduction ratio comprised between 3 and 4.2
 4. The gas turbine engine ofclaim 1, wherein the pivot axis extends along a radial direction.
 5. Thegas turbine engine of claim 1, wherein the fan has a diameter in a rangefrom 220 cm to 380 cm
 6. The gas turbine engine of claim 1, wherein thefan has a diameter in a range from 220 cm to 300 cm and is configured tohave a rotational speed at cruise conditions in a range from 1700 rpm to2500 rpm.
 7. The gas turbine engine of claim 1, wherein the fan has adiameter in a range from 330 cm to 380 cm and is configured to have arotational speed at cruise conditions in a range from 1200 rpm to 2000rpm.
 8. The gas turbine engine of claim 1, wherein a bypass ratiodefined as a ratio of mass flow rate of the bypass airflow to mass flowrate of the core airflow at cruise conditions is in a range from 10 to12.
 9. The gas turbine engine of claim 1, wherein a bypass ratio definedas a ratio of mass flow rate of the bypass airflow to mass flow rate ofthe core airflow at cruise conditions is in a range from 13 to
 15. 10.The gas turbine engine of claim 1, wherein the compressor is a firstcompressor, the turbine is a first turbine, and the shaft is a firstshaft, the engine core further comprising a second compressor downstreamof the first compressor, a second turbine upstream of the first turbine,and a second shaft connecting the second turbine with the secondcompressor.
 11. A gas turbine engine comprising: a fan rotating about anengine main axis and generating a core airflow and a bypass airflow; acore duct, across which the core airflow flows; an engine corecomprising: a compressor for compressing the core airflow and comprisinga plurality of stages, each stage comprising a row of rotor blades and arow of stator vanes, a first stage of said plurality of stages beingarranged at an inlet of the compressor; combustion equipment; and aturbine connected to the compressor through a shaft; an Engine SectionStator (ESS) comprising a plurality of ESS vanes and arranged in thecore duct downstream of the fan, each ESS vane comprising an ESS leadingedge and an ESS trailing edge; and a plurality of variable inlet guidevanes (VIGVs) adapted to rotate about a pivot axis and arranged in thecore duct downstream of the ESS and upstream of the compressor, eachvariable inlet guide vanes (VIGVs) comprising a VIGV leading edge and aVIGV trailing edge, and wherein the plurality of VIGVs are arrangedangularly rotated with respect to the plurality of ESS vanes such thatfirst longitudinal planes passing through respective 70% span ESSleading edge points are angularly rotated with respect to correspondingsecond longitudinal planes passing through respective 70% span VIGVpivot axis points by a rotation angle α, and wherein the plurality ofVIGVs are arranged angularly rotated with respect to the plurality ofESS vanes such that longitudinal planes passing through respective 90%span ESS leading edge points are angularly rotated with respect tocorresponding longitudinal planes passing through respective 90% spanVIGV pivot axis points by a rotation angle α1 less than the rotationangle α at 70% span and comprised between 0.05° and 5°.
 12. The gasturbine engine of claim 11, wherein a mid-span ESS leading edge point isat an axial distance L from a mid-span VIGV leading edge point, thedistance L being comprised between 80 mm and 650 mm.
 13. The gas turbineengine of claim 11, further comprising a strut arranged in the core ductbetween the ESS and the plurality of VIGVs.
 14. The gas turbine engineof claim 11, wherein the ESS comprises 40 to 80 ESS vanes
 15. A gasturbine engine comprising: a fan rotating about an engine main axis andgenerating a core airflow and a bypass airflow; a core duct, acrosswhich the core airflow flows; an engine core comprising: a compressorfor compressing the core airflow and comprising a plurality of stages,each stage comprising a row of rotor blades and a row of stator vanes, afirst stage of said plurality of stages being arranged at an inlet ofthe compressor; combustion equipment; and a turbine connected to thecompressor through a shaft; an Engine Section Stator (ESS) comprising aplurality of ESS vanes and arranged in the core duct downstream of thefan, each ESS vane comprising an ESS leading edge and an ESS trailingedge; and a plurality of variable inlet guide vanes (VIGVs) adapted torotate about a pivot axis and arranged in the core duct downstream ofthe ESS and upstream of the compressor, each variable inlet guide vanes(VIGVs) comprising a VIGV leading edge and a VIGV trailing edge, andwherein the plurality of VIGVs are arranged angularly rotated withrespect to the plurality of ESS vanes such that first longitudinalplanes passing through respective 70% span ESS leading edge points areangularly rotated with respect to corresponding second longitudinalplanes passing through respective 70% span VIGV pivot axis points by arotation angle α, and wherein the plurality of VIGVs are arrangedangularly rotated with respect to the plurality of ESS vanes such thatlongitudinal planes passing through respective 90% span ESS leading edgepoints are angularly rotated with respect to corresponding longitudinalplanes passing through respective 90% span VIGV pivot axis points by arotation angle α1 greater than the rotation angle α at 70% span andcomprised between 0.02° and 6°.
 16. The gas turbine engine of claim 15,further comprising a reduction gearbox that receives an input from theshaft and outputs drive to the fan so as to drive the fan at a lowerrotational speed than the shaft.
 17. The gas turbine engine of claim 16,wherein the reduction gearbox has a gear ratio in a range from 3.1 to3.8.
 18. The gas turbine engine of claim 15, wherein the engine isconfigured to have a Turbine Entry Temperature (TET) at cruiseconditions in a range from 1400 K and 1650 K.
 19. The gas turbine engineof claim 15, wherein the engine is configured to have a Turbine EntryTemperature (TET) at maximum take-off condition in a range from 1800 Kand 1950 K.
 20. The gas turbine engine of claim 15, wherein a mid-spanESS leading edge point is arranged at a first radius (R1) from theengine main axis and a mid-span VIGV leading edge point is arranged at asecond radius (R2) from the engine main axis, wherein a difference ΔRbetween the first radius and the second radius is between 60 mm and 280mm.